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A short history of the European Transonic Wind Tunnel ETW

John Green a,n, Jurgen Quest b

aAircraft Research Association, Manton Lane, Bedford MK41 7PF, United Kingdomb ETW GmbH, Ernst-Mach-Strasse, 51147 Koln, Germany

a r t i c l e i n f o

Keywords:

Wind tunnel history

Cryogenic wind tunnels

Wind tunnel design

Wind tunnel test techniques and model

instrumentation

Transonic aerodynamics

European aeronautical collaboration

a b s t r a c t

This paper is written as a contribution to the celebration of 50 years of Progress in Aerospace Sciences

and of the centenary of the birth of its founder, Dietrich Kuchemann. It reviews the evolution of the

European Transonic Wind Tunnel, ETW, from early conceptual studies to its entry into service and its

current capabilities and achievements. It traces the development, from the earliest days, of experi-

mental aerodynamics and of the basic aerodynamic understanding that gave rise to the main periods of

wind tunnel building before and after World War II. By about 1960, this activity appeared to have come

to a natural halt. The paper gives an account of the role of Kuchemann in arguing the need in 1968 for a

further step in wind tunnel capability, to provide transonic testing at high Reynolds numbers. It

describes his leading role in gaining acceptance of the concept, formulating the specification and

promoting studies of alternative, radical design options for the co-operative European project that

became ETW. The progress of ETW through design, construction, commissioning and into full operation

is recorded. The paper discusses the many technical innovations that have been introduced in order to

meet customer requirements in the challenging field of aerodynamic testing in a cryogenic environ-

ment and, finally, looks to the future and the further technical challenges that it holds.

& 2011 Elsevier Ltd. All rights reserved.

Contents

1. Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 320

2. Experimental aerodynamics in the beginning 17421917. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 321

2.1. Early insights, 17421904 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 321

2.2. The evolution of the wind tunnel up to 1917. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 323

3. The coming of age of the wind tunnel 19171945 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 325

3.1. The pursuit of full scale Reynolds numbers in the 1920s and 1930s. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 325

3.2. The significance of compressibility and the first high-speed tunnels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 326

4. The great period of wind-tunnel building 19451959 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 327

5. Emergence of the need for higher Reynolds number 19591968 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 329

6. Definition of the requirement and the solution for Europe 19681978 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 331

6.1. The role of AGARD and Kuchemann . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 331

6.2. The work of LaWs and MiniLaWs 19711974 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 333

6.3. AEROTEST and AC/243 (PG.7) 19721973. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3346.4. The LaWs specification and the four original design concepts. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 335

6.4.1. The transonic Ludwieg Tube tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 336

6.4.2. The Evans Clean Tunnel (ECT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 337

6.4.3. The Injector-Driven Tunnel (IDT). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 338

6.4.4. The Hydraulic-Driven Tunnel (HDT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 338

6.5. Engineering studies of the four design concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 339

6.6. The coming of cryogenics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 341

6.7. The underlying physics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 343

6.8. Evolution of the specification, from LEHRT to ETW 19751978 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 345

Contents lists available at ScienceDirect

journal homepage: ww w.elsevier.com/locate/paerosci

Progress in Aerospace Sciences

0376-0421/$- see front matter& 2011 Elsevier Ltd. All rights reserved.

doi:10.1016/j.paerosci.2011.06.002

n Corresponding author. Tel./fax: 44 1525 290631.

E-mail address: [email protected] (J. Green).

Progress in Aerospace Sciences 47 (2011) 319368
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7. Designing the ETW 19781988 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 346

7.1. Phase 2.1 preliminary design 19781985 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 346

7.2. Phase 2.2 final design and the Rogers task force 19851988. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 351

8. Establishing the GmbH 1988. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 352

9. The construction phase 19881993 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 352

10. Hardware characteristics of the facility . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353

10.1. The settling chamber and its downstream contraction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353

10.2. The drive system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 354

10.3. The nitrogen system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 354

10.4. Model handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35410.5. The test section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 355

11. Getting Wind on 19932000 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 356

11.1. Tuning and calibrating. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 356

11.2. Client testing in the 1990s. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 357

11.3. Developing techniques for gathering fully corrected data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 357

12. Operating in the 21th century 20002010 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 359

12.1. Contributing to European research . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 359

13. Developing and enhancing test techniques. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 360

13.1. Further development and enhancement of test techniques . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 363

13.2. Laminar wings are back. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 365

14. Summary and outlook . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366

References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 367

Further reading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 368

1. Introduction

The website of ETW GmbH boldly asserts, The European Transo-

nic Wind Tunnel, ETW, in Cologne, Germany, is the most modern

wind tunnel in the world; a unique test facility for the development

of new transport aircraft. Fig. 1is an aerial view of the facility. This

paper gives an account of its evolution, achievements to date, current

capabilities and the part that Dietrich Kuchemann played in its

creation.

There are two high Reynolds number transonic wind tunnels

in the world, the National Transonic Facility (NTF) at NASA

Langley and ETW in Cologne. The tunnels are similar in size and

operating principle both are cryogenic, using gaseous nitrogen

at near liquefaction temperatures as the working fluid and bothfar surpass all other wind tunnels in their ability to test aircraft

models at Reynolds numbers equal to, or near to, those of flight.

The NTF first ran in 1983. Its reported cost was $85 million. ETW

first ran 10 years later, in 1993. Its construction cost was 562

million Deutschmark at 1987 prices, roughly twice the cost of the

NTF when adjusted for inflation, and in some key respects it is a

more advanced facility than the NTF. But we recall that Isaac

Newton said, If I have seen further, it is only by standing on the

shoulders of giants. ETW and Isaac Newton may seem an unusual

juxtaposition, but there is a parallel here, in that the ETW has

beyond doubt benefited immensely from the pioneering work of

NASA that led to the NTF. Without the NTF, there would be no

ETW as we know it today.

It is also possible that, without Dietrich Kuchemann, there

would be no ETW at all. It was he who led the intellectual debate

within NATO that resulted eventually in four nations, France,

Germany, The Netherlands and the United Kingdom, deciding in

1973 to co-operate in a project to build a high Reynolds number

transonic tunnel for Europe. The drive towards such a tunnel was

triggered by events in the 1960s that were reported at a Specialists

Meeting[1]of the AGARD Fluid Dynamics Panel (FDP) in September

1968. The subject of the meeting was Transonic Aerodynamics andKuchemann, who was a member of the Programme Committee, was

asked to prepare a Technical Evaluation Report on the meeting.

It was a happy chance for Europe that the task fell to Kuchemann.

He was, at that time, the Head of Aerodynamics Department at

the Royal Aircraft Establishment in Farnborough, internationally

respected and with deep insight into the application of the results

of aerodynamic research to aircraft design, particularly the design of

aircraft operating at transonic conditions. He was also a believer in

getting things done rather than merely philosophising and his

energy and commitment to making progress played a vital part in

shaping the concept of a co-operative European transonic tunnel

and in convincing the four nations of the need for it.

Much of the work that was done under his leadership was on

alternative, novel concepts for a tunnel with air at ambienttemperature as the working fluid. By 1974, however, the concept

of a cryogenic transonic tunnel, in which higher Reynolds num-

bers are achieved by testing in gaseous nitrogen at very low

temperatures, had been shown at NASA Langley to be an attrac-

tive possibility. In October 1975, at the Specialists Meeting of the

AGARD FDP on Wind Tunnel Design and Testing Techniques [2],

the first paper was from NASA Langley, presenting the results

obtained in a pilot cryogenic transonic tunnel and setting out

the plans for the NTF. Kuchemann, in his closing address as the

outgoing Chairman of the Panel, noted the potential of the

cryogenic tunnel, welcomed the progress that had been made in

the USA and, speaking of Europe, ended his address, So all I need

to do now is to quote one of the speakers who said: Now, let us do

it! Just over 4 months later, on 23 February 1976, he diedFig. 1. Aerial view of the European Transonic Wind Tunnel.

J. Green, J. Quest / Progress in Aerospace Sciences 47 (2011) 319 368320


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unexpectedly, after a short illness. He was not to see his vision

come to fruition but, through his inspirational leadership, he had

set the course that the four nations followed to its logical

conclusion. There were obstacles and delays on the road ahead

but the end result meets the requirements that were set out

under his guidance in the early 1970s and stands as a testament

to his insight, perseverance and ability to inspire his colleagues.

The authors of this paper span more than 40 years of involvement

in ETW. In 1970 John Green, a member of Aerodynamics Departmentat the Royal Aircraft Establishment, was invited (instructed) by his

Department Head (Kuchemann) to write a paper on viscous flows

over wings for the AGARD FDP meeting[3]in May 1971, for which

Kuchemann was Chairman of the Programme Committee. In the

following years he contributed other papers to the studies led by

Kuchemann. In the period 19781981, the early years of Phase 2.1,

the Preliminary Design phase of the ETW project, he was the UK

member of the Steering Committee,1 chairing the committee in

1980. In the early 1990s, having left Government service, he was a

member of ETW Advisory Committee 1.2 He has written the first part

of the paper, covering the years from 1742, when projectile drag at

transonic speeds was first measured, to 1988 when the four nations

resolved to carry the ETW project through to its conclusion. Jurgen

Quest joined ETW in 1988, after the Technical Group had moved

from Amsterdam to Cologne, at the time of the establishment of ETW

GmbH and the start of Phase 3, the construction and operation phase.

He is currently ETW Chief Aerodynamicist and he has written the

second part of the paper, from 1988 to the present day.

2. Experimental aerodynamics in the beginning 17421917

2.1. Early insights, 17421904

In the late 19th century and for the first third of the 20th century,

the University of Gottingen was an academic centre of world

renown.3 It was host to an outstanding collection of mathematicians

and scientists whose research in many fields, not least aerody-

namics, paved the way for many of the advances of the 20thcentury. Dietrich Kuchemann was born and educated in Gottingen,

took his doctorate at Gottingen University under Ludwig Prandtl in

1936 and, for the next 10 years, continued his aerodynamic research

at the A.V.A. (Aerodynamische Versusch Anstalt) Gottingen. There is

no doubt that he was imbued with the spirit that prevailed there in

his student years. In May 1975 he spoke eloquently of the Gottingen

spirit der Gottinger Geist in his opening address to a symposium

of the AGARD Fluid Dynamics Panel on Flow Separation held in the

town[4]. He characterised the spirit as a determination to know, to

understand, coupled with the firm intention that knowledge gained

should be applied usefully, should be of benefit to human society.

That spirit was clearly in evidence throughout his determined efforts

to drive forward the European studies that led finally to the ETW.

We begin this paper with a brief review of the evolution of ourunderstanding of the aerodynamic phenomena that gave rise to the

need for Europe to build the ETW. Central to that understanding are

three fundamental conceptual advances, all linked in some way to

Gottingen but, to begin, we go back to the 18th century, and to

experiments rather than theory, for the starting point in our narrative.

In 1742 Benjamin Robins devised a ballistic pendulum, Fig. 2,which he employed to make what were, for that time, some

remarkably accurate measurements of the drag of a ball fired

from a musket [5]. His purpose was to demonstrate that the

resistance of the air had an important influence on the trajectory

of a cannon ball and that, as a consequence, all ballistic calcula-

tions at that time, which took the resistance to be negligible, were

ill founded. In this he succeeded, but he also discovered that, over

the range of velocities covered by his experiments, the drag of the

ball did not vary as the square of the velocity as predicted by

the accepted authority at that time, Sir Isaac Newton [6]. Over the

speed range that his experiments covered, (Mach 0.71.5 in

modern terminology) he found that the drag increased more

rapidly than the square of the velocity (Fig. 3). He had measured

transonic drag rise[7].Ernst Mach, born almost 100 years after Robins published his

paper, also studied ballistics experimentally. His most notable

contribution was the use of Schlieren photography to observe

gunshots and to display the pattern of shock waves created by a

bullet travelling at high speed. He observed that the inclination

of the shock wave was a function of the ratio of the speed of

the bullet to the speed of sound. This observation was initially of

little interest to aeronautical scientists until, as flight speeds

increased, one of the leaders in the field, Jakob Ackeret, who

had worked under Prandtl in Gottingen from 1921 to 1927,

published in 1927 an article on gasdynamics [8] in which he

proposed the term Mach number, MV=a, for this ratio of

velocities. Flight Mach number is the first parameter that ETW

is required to replicate.

Fig. 2. Ballistic pendulum of Benjamin Robins, 1742.

Source: Ackroyd, UKs contribution to development of aeronautics, Part 1, Aero J.,

January 2000.

1 With the foundation of ETW in 1988, the Steering Committee was expanded

slightly and became the Supervisory Board, the governing body of ETW.2 AC 1 was established by the Supervisory Board to provide advice on matters

related to the expected development of aerospace science and engineering,

especially in Europe.3 The coming to power of the Nazi party in Germany in 1933 was followed

almost immediately by the great purge of Jewish scientists, which resulted in

many of the most distinguished academics at Gottingen leaving the country.

J. Green, J. Quest / Progress in Aerospace Sciences 47 (2011) 319 368 321


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Shortly after Ernst Machs experiments in Prague with gunshots,

Osborne Reynolds, in Manchester in 1883, made some rather different

but no less significant experiments with water flowing through a

tube. The tube was of glass and a thin stream of coloured water

flowing down the middle of the tube was used to visualise the

behaviour of the flow. The man, his apparatus and the flow patterns

that he observed are shown inFig. 4. Patterns ac, seen by spark

illumination at three different, increasing flow rates, were termed by

Reynolds direct for a and sinuous for c today we call this laminar

and turbulent flow. By performing experiments with tubes of three

different diameters, and by varying the temperature of the water

and the flow rate through the tube, Reynolds established that the

character of the flow depended on a non-dimensional quantity

RnrVl=mwherer is the fluid density,Vits velocity,m its absoluteviscosity andl a length scale (in his experiments Reynolds chose the

tube diameter). At low values of this quantity the flow was laminar,at high values turbulent. From a consideration of the equations of

motion, Reynolds reasoned that the quantity represented the ratio of

inertial to viscous forces acting on a small volume of fluid and that

at some characteristic value of this quantity the flow would begin to

form eddies.

The term Reynolds number for the quantity Rn4 was first

proposed in 1908 [9] by the physicist Arnold Sommerfeld, a

graduate of Gottingen. But, even before it had been given a name,

its fundamental importance to aerodynamics had been recognised

and both Lord Rayleigh[10], in his 1884 Presidential Address to

the British Association in Montreal, and Lanchester [11]in 1907,

in his seminal book Aerodynamics, had identified equality of

this quantity as a requirement for fluid flows to be dynamically

similar.

In the early years of flight, although the significance of Reynolds

number was recognised, it was understood that it could not be

replicated in the ground test facilities of the time, whirling arms and

small wind tunnels. The true full-scale aerodynamics could be

realised only in flight by the full-scale machine. Fortunately, the

aerodynamic properties of early aircraft were not strongly depen-

dent on Reynolds number and failure to replicate flight values in the

ground test facilities of the time did not seriously undermine the

usefulness of these facilities.

In the period immediately after World War II, when many new,

large wind tunnels were built, both Mach number and Reynolds

number were recognised as important parameters of the new

generation of high speed aircraft. Although it was now possible to

replicate flight Mach numbers in the wind tunnel, the maximum

achievable Reynolds numbers were lower than flight by an order ofmagnitude. Hence the post-war practice evolved of testing and

reporting results at specific Mach numbers and of developing

methods of adjusting the data for the difference in Reynolds number

between tunnel and flight. This approach appeared to be satisfactory

for the first two decades that followed World War II.

Eventually, however, the approach was undermined by advances

in wing design that increased the importance of the behaviour of the

wing boundary layer. It was in 1904 that Prandtl, then a professor of

mechanics at the technical school in Hannover, presented his theory

of the boundary layer at the Third International Mathematical

Congress at Heidelberg[12]. Its impact was great and was a factor

no doubt in his appointment as director of the Institute for Technical

Physics at the University of Gottingen later that year. In fact,

Prandtls theory enabled the gulf between the theoretical results of19th century hydrodynamicists and the practical results of the

aeronautical experimenters finally to be bridged. It is arguably the

single most important concept in the evolution of aerodynamics,

explaining the key role of the boundary layer in determining

aerodynamic behaviour and also enabling the full power of inviscid

flow theory to be brought to bear on predicting the flow about

aircraft. In the years after 1904, Prandtl and his colleagues in

Gottingen made many further, fundamentally important contribu-

tions to our understanding of the flow about aerofoils and wings and

the behaviour of the boundary layer.

Fig. 3. Comparison between sphere drag measured by Robins using ballistic pendu-

lum and present day result.

Source:Ref.[7],Fig. 9.

Fig. 4. Osborne Reynolds with experimental apparatus and observed flow

patterns, 1883.

Source:internet.

4 The usual symbol for Reynolds number is R. Here we useRn so that we may

useR for the universal gas constant.



5/50

Theory and experimental capabilities advanced together in the

early decades of the 20th century with the support of wind tunnel

testing becoming increasingly important in the development of any

new aircraft. As aircraft design evolved and flight speeds increased,

so wind tunnel facilities grew in size, complexity and cost. By 1904,

the three key concepts, Mach number, Reynolds number and the

boundary layer had emerged. However, more than 60 years were to

pass before the need for a wind tunnel to simulate not only Mach

number but also to approximate flight Reynolds number closely, inorder to simulate the behaviour of the boundary layer in flight as

accurately as possible, became generally recognised.

2.2. The evolution of the wind tunnel up to 1917

In 1742, Benjamin Robins determined the drag of musket balls

by measuring the velocities of balls fired over different ranges

with a fixed charge of powder and calculating the deceleration

from the reduction in impact velocity with increase in range. Four

years later, in 1746, he reported experiments with a whirling arm

apparatus (Fig. 5) in which a weight rotated a drum that carried

the test object on a long arm. Drag was determined by the weight

while velocity of the test object was measured by timing a

number of revolutions of the arm. This gave him more accuratedrag data, for a range of shapes, but only in low speed flow.

The whirling arm concept was taken up by others, notably

Sir George Cayley, who in 1804 used a whirling arm to measure

the lift force on a square plate at angles of incidence between

31 and 181 (Fig. 6). Using these data he designed, built and

successfully flew a model glider (Fig. 7) believed to have been

the first successful heavier than air vehicle in history. In the 19th

century several other researchers used the whirling arm, notably

Otto Lilienthal, who between 1866 and 1889 built several whir-

ling arms of different sizes and measured the lift and drag

characteristics of a variety of aerofoils[13]. He also made similar

measurements of the forces on stationary aerofoils in the wind

over open ground. Because the whirling arm created a swirling

motion in the air around it, there were doubts about the validity

of the data it produced and Lilienthal concluded that his mea-

surements in the natural wind were the more reliable. He used

these in the design of the gliders in which he made more than

2500 flights between 1891 and his final, fatal flight in 1896. In

1895 he published tables derived from his natural-wind measure-

ments and these, republished in the USA in 1897, were used by

the Wright brothers to design their gliders of 1900 and 1901.

Meanwhile, in Britain, Francis Wenham, following unsatisfactory

experiments with a whirling arm, in 1871 persuaded the Aeronau-

tical Society of Great Britain to raise the funds to build a wind

tunnel, the worlds first. It consisted of a duct 12 ft long and

18 in18 in in cross section with a fan upstream of the model

driven by a steam engine. It had poor flow quality but nevertheless,

from tests on a variety of wing shapes, two significant results

emerged. First, that at small angles of incidence the lift force varies

in proportion to the sine of the angle of incidence, rather than to the

square of the sine. Secondly, that wings of high aspect ratio had

higher lift to drag ratios than those of low aspect ratio.5 In the early

1880s, also in Britain, Horatio Phillips built a wind tunnel of similar

proportions but driven by a steam ejector. This produced a steadierflow and led to Phillips developing and patenting a series of

cambered aerofoils, considered the first truly modern aerofoils.

Others followed Wenham and Phillips in building and experiment-

ing in wind tunnels, but with little further impact until the decisive

step forward taken by the Wright brothers in the autumn of 1901.

The Wrights had designed their first glider using Lilienthals

tables of normal and axial force. When they took it to Kitty Hawk,

North Carolina in September 1900, they had some limited success

but found that its lift was rather lower than had been expected.

Results the following year, with a new glider with increased wing

area, also fell well below expectations. The Wrights concluded that

Lilienthals tables were not reliable6 and in the autumn of 1901 built

themselves a wind tunnel similar to Wenhams, with a 16 in16 in

test section and a two-bladed fan driven by a gasoline engine(Fig. 8). They measured the lift and drag of some 200 model wings

Fig. 5. Benjamin Robins whirling arm, 1746.

Source:NASA Centennial of Flight.www.centennialofflight.gov.This is a re-drawn

version of the Robins original. The latter is in Ackroyd on the same page asFig. 2.

Fig. 7. Sir George Cayleys glider, 1804.

Source: Scanned from download from rsnr.royalsocietypublishing.org, paper by

Ackroyd on Cayley (2002) p. 175.

Fig. 6. Sir George Cayleys whirling arm, 1804.

Source:rsnr.royalsocietypublishing.org, paper by Ackroyd on Cayley (2002) p. 173.

5 The sine squared law, Newtons theory [6], had led to the widely accepted

conclusion that heavier than air flight was not practicable. Others, notably Cayley,

had found a linear variation of lift with incidence for low aspect ratio surfaces but

it was Wenhams discovery of high lift to drag ratios for high aspect ratio surfaces

that gave members of the Aeronautical Society reason to believe that heavier than

air flight would one day be achieved.6 The error in Lilienthals tables, which were based on his measurements of

the forces on an aerofoil in a natural wind, arose from his use of a plate

anemometer calibrated on a value of plate drag coefficient quoted by Smeaton

in 1759 from whirling arm results obtained by his friend, a certain Mr. Rouse of

Harborough. The Wright Brothers determined from their wind tunnel tests that

the Smeaton coefficient was incorrect and should have been 0.0033 rather than

the 0.005 that had been widely used for the previous century and a half.

http://www.centennialofflight.gov/http://www.centennialofflight.gov/http://localhost/var/www/apps/conversion/tmp/scratch_3/rsnr.royalsocietypublishing.orghttp://localhost/var/www/apps/conversion/tmp/scratch_3/rsnr.royalsocietypublishing.orghttp://www.centennialofflight.gov/


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with different aerofoil sections and planform, using a simple balance

of their own design which gave accurate and repeatable results.

Their 1902 glider (Fig. 9) designed on the basis of their wind tunnel

results, had nearly twice the span of the 1900 glider. At Kitty Hawk,

in 5 weeks in September and October during which they made

between 700 and 1000 glides, the brothers developed the flight

controls for this machine so that it was fully controllable in three

dimensions. It was also more efficient aerodynamically than any-

thing that had gone before, with a lift-to-drag ratio of 8. They had

established a solid basis for the larger machine with which, inDecember 1903, they made the first controlled powered flights. It is

an achievement that would not have been possible without their

wind tunnel test programme in 1901.

Early in the 20th century, Gustav Eiffel began aerodynamic

investigation by measuring the aerodynamic forces on objects

dropped from the second platform of the Eiffel Tower, 377 ft above

ground level. He followed this in 1909 by building a wind tunnel of

novel design in the shadow of the Eiffel Tower. Its test section was

1.5 m diameter and its fan was driven by an electric motor drawing

on the towers power supply. In 1912 he built a similar but larger

tunnel at Auteuil (Fig. 10) and patented the design. Like the earlier,

smaller tunnels of Wenham, Phillips and the Wrights, it was an open-

return tunnel, housed in a hangar, but having an open jet test section

with air drawn into the jet nozzle though a bellmouth by a fan at the

outlet from the diffuser downstream of the test section. Eiffels

introduction of the bellmouth and diffuser meant that the pressure

in the test section was lower than the pressure in the hangar and the

test section therefore had to be inside a hermetically sealed enclo-

sure, the experimental room. Eiffels experiments led to a number of

significant advances; he pioneered the testing of models of complete

aircraft and, in resolving a factor of two disagreement between his

results and those of Prandtl in Gottingen, in 1914 he demonstrated

for the first time the sharp drop in the drag of a sphere as Reynolds

number is increased above 300,000 approximately, when the

boundary layer on the sphere changes from laminar to turbulent

[13]. The Eiffel wind tunnel concept was considered a success and

further, larger versions were built in the following decades.

Some 600 km to the North-East in Gottingen, at Prandtls

suggestion, the German Society for Airship Study (Motorluftschiff

Studiengesellschaft) in 1907 funded the construction of a simple

wind tunnel at a cost of 20,000 marks. It had a closed returncircuit of rectangular planform with a closed test section 2 m2.

There was a honeycomb flow straightener downstream of the fan

in the return leg and cascades of turning vanes at each corner.

However, because almost the entire circuit had the same cross-

sectional area and flow velocity as the test section, flow quality in

the test section was not particularly good. The tunnel was

constructed in 1908 and in 1909 began practical work on the

aerodynamics of airships. It was envisaged as a temporary facility

and in 1911 Prandtl made the first case for building something

more substantial. Negotiations for the funds for this were essen-

tially complete in 1914 when World War I broke out, the plans

were put on hold and the first wind tunnel, now concentrating on

aircraft aerodynamics, continued as an important test facility for

most of the war. In 1915 the case to build a second tunnel wasaccepted by the war administration and 300,000 marks were

made available15 times the funding for the first tunnel. The

project was completed and began operations in Spring 1917.

The second Gottingen tunnel (Fig. 11) was a great advance on

what had gone before and embodied for the first time many features

that have become standard in most tunnels built since then. In fact,

in the years that followed, wind tunnels tended to be classed as

either the Eiffel or the Gottingen design. The key features of the

Gottingen design were explained by Prandtl in a lecture in 1920

[14]. He had combined the idea of a contraction ahead of the open-

jet test section and a diffuser downstream, a concept he acknowl-

edged as coming from Eiffel, with a closed return circuit of

substantially greater cross-sectional area, and hence lower flow

velocities, than the test section. The Eiffel contraction and diffuser

Fig. 9. Wilbur Wright in the 1902 glider.

Source: NASA Centennial of Flight: www.centennialofflight.gov.

Fig. 10. Eiffel wind tunnel at Auteuil, 1912.

Source:The wind tunnels of NASA, NASA SP-440 Chapter 2.

Fig. 8. Wright Brothers wind tunnel, 1901.

Source:www.wright-brothers.org.

http://www.centennialofflight.gov/http://www.centennialofflight.gov/http://www.wright-brothers.org/http://www.wright-brothers.org/http://www.centennialofflight.gov/


7/50

increased the efficiency of the circuit and reduced the fan power

requirement while the closed return circuit, ending in a settling

chamber with a flow smoothing honeycomb, followed by a 5:1

contraction, produced a more uniform and steady flow than in

previous tunnels. The closed return circuit removed the need for the

testing room to be hermetically sealed, as was the case in the Eiffel

tunnel, and hence made the test area more accessible. Models were

mounted on carts that ran on transverse tracks and there were bays

on either side of the test section to enable one model to be prepared

while another was in the test sectiona feature that Prandtl

considered important and that has since been replicated in a

number of major wind tunnels, including ETW. The fan was driven

by a Ward Leonard set, an AC motor driving a DC generator to

supply the DC motor driving the fan. This gave accurate speed

control over a range from 50 to 1100 rpm and set a pattern that

became generally adopted for subsonic wind tunnel drive systems.

The use of testing with the model erect and inverted to determine

with precision the inclination of the flow in the tunnel, relative to

the direction of gravity, was another innovation that remains best

practice in todays wind tunnels.

There were some innovations, such as building the tunnel out of

reinforced concrete with its circuit in a vertical plane, that have beencopied less frequently, some of the advances in measurement

techniques and automatic speed control have been superseded

and, increasingly, closed test sections have been preferred to open

jets. Even so, many of the key features of the Gottingen tunnel can

be found in almost all the worlds major wind tunnels built since

1920. And two other practical considerations noted by Prandtl in

1920 have featured in the building of many major wind tunnels

since. First, though the drive power of the Gottingen tunnel was only

about 0.25 MW, that power was significant relative to the capacity

of the Gottingen local power supply at the time and an automatic

regulator was needed to avoid making large load increases suddenly.

Power availability has since featured in decisions on the location of a

number of large wind tunnels. Secondly, the design of the tunnel

circuit strikes a balance between running and capital costs. Thetunnel circuit is shorter and less aerodynamically efficient than

optimum, thereby reducing the cost of the tunnel shell. This trade-

off between capital and running costs has to be made in the design

of every major wind tunnel and was an important consideration

during the assessment of alternative drive systems for ETW.

3. The coming of age of the wind tunnel 19171945

3.1. The pursuit of full scale Reynolds numbers in the

1920s and 1930s

At the end of World War I the Gottingen tunnel could be

considered the state of the art. If we adopt the convention used

in specifying the ETW Reynolds number, that a typical wing chord is

0.1 times the square root of the cross-sectional area of the test

section, the maximum Reynolds number of the Gottingen tunnel

based on this typical chord was approximately 0.7 million. For the

Sopwith Camel, a typical WWI fighter aircraft, the chord Reynolds

number was approximately 4.7 million. For a larger aircraft, such as

the Vickers Vimy bomber, it was approximately 9 million. There was

thus an order of magnitude difference between characteristic tunnel

and flight Reynolds numbers.

Despite the Wright brothers having made the first controlled

powered flight in 1903 and having taken Europe by storm with

Wilbur Wrights demonstration of the abilities of the Wright Flyer in

Paris in 1908, aeronautical progress in the USA lagged far behind

progress in Europe in the following decade. This was recognised in

the USA as early as 1912 but it was not until March 1915 that

Congress, at the recommendation of the regents of the Smithsonian

Institution, passed legislation to establish the National Advisory

Committee for Aeronautics (NACA). In 1917 NACA established a

laboratory site in Hampton, Virginia and named it Langley Field.

NACA Wind Tunnel No. 1, a low speed tunnel of the Eiffel type with

a test section 1.5 m in diameter, began operation at Langley Field in

June 1920. Its characteristic Reynolds number based on one tenththe square root of its test section area was 0.37 million; its life was

relatively short and unproductive.

A year later, in June 1921, the NACA Executive Committee

decided to build a much more substantial and important tunnel,

the Langley Variable Density Tunnel (VDT), in order to test at

Reynolds numbers much closer to full scale flight. With Reynolds

number defined as

RnrVlm

,

wherer is the density,Vthe velocity,l a characteristic length and mabsolute viscosity, Rn can be increased, in air at ambient tempera-

ture, by increasing pressure and thereby density, increasing velocity

or increasing the characteristic length. The concept of increasing itby increasing pressure was proposed by Max Munk, who had

obtained his doctorate in Gottingen under Prandtl and had moved

to the USA to a post in NACA Headquarters in Washington in 1920.

Shown inFig. 12, the tunnel had a circular cross section with a

closed test section of 5 ft (1.5 m) diameter followed by a diffuser

embedded within an annular return circuit, all contained within a

cylindrical pressure vessel. The maximum velocity in the test

section was 50 mph, as against 90 mph for the no. 1 wind tunnel,

but it could test at pressures up to 20 atm and thus had a

characteristic Reynolds number of 4.2 million, comparable to

the chord Reynolds number of fighter aircraft of the time. The

tunnel became operational in 1923 and was used to obtain high

Reynolds number data on a wide range of aircraft and airship

types, A particularly substantial contribution was its testing of

Fig. 11. Prandtls wind tunnel at Gottingen, 1916.

Source:Ref. [13]p. 300.



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aerofoil sections; the aerodynamic data on 78 sections, published

in 1933 in NACA Technical Report 460, was an important land-

mark which prepared the ground for the design of the many new

aircraft developed in the USA before and during World War II.

The UK was impressed by the results from the VDT and, a

decade after the USA, built a similar tunnel at the National

Physical Laboratory (NPL) at Teddington. Over that decade Jones

[15] had published his paper on The Streamline Aeroplane,

aircraft design had progressed from biplanes braced with struts

and wires to the streamlined monoplanes competing in the

Schneider Trophy, the world airspeed record had doubled and,

with it, flight Reynolds numbers. Accordingly the NPL Com-

pressed Air Tunnel (CAT), which had essentially the same layout

as the NACA VDT, had a test section diameter and maximum

speed both 20% greater than the VDT. These increases, together

with an increase in maximum pressure to 25 atm, gave the CAT a

characteristic Reynolds number of 7.5 million.

Although the VDT and CAT were valuable sources of high

Reynolds number data, their layout, with flow smoothing honey-

combs in the relatively high speed stream immediately upstream

of the test section, resulted in comparatively high free stream

turbulence levels. These, as Doetsch [16] showed in 1936 fromtunnel turbulence data and NACA deduced from comparisons

between wind tunnel and flight, adversely affected the wind

tunnel results. The turbulence was understood to promote pre-

mature transition, a particular problem in research to develop

laminar flow aerofoils, and thus low turbulence became a desir-

able feature for future tunnels.7 To explore this question, NACA

built a pilot low turbulence tunnel which came into operation in

1939, designated the NACA Ice Tunnel.

NACA used the Ice Tunnel as the basis for the design of the Low

Turbulence Pressure Tunnel (LTPT) which went into operation at

Langley in 1941. The tunnel had a contraction ratio of 17.6:1, with

a combination of gauze screens and a honeycomb in the settling

chamber to minimise test section turbulence. The test section was

7.5 ft high3 ft wide, intended specifically for two-dimensionalaerofoil testing. With a maximum pressure of 10 atm and a max-

imum speed which varied with tunnel pressure but was about

130 mph at maximum pressure; its characteristic Reynolds number

was about 5.8 million. However, for tests on a two dimensional

aerofoil that fully spanned the tunnel, a chord of 2.0 ft was normal

and tests were typically done at Reynolds numbers of 3.0, 6.0 and

9.0 million. The tunnel played a key role in the development of the

NACA 6-series of laminar flow, low drag aerofoils that were adopted

for later WWII aircraft such as the highly successful P-51 escort

fighter.

Sixteen years before the LTPT went into service, and only 2

years after the VDT began operations, NACA decided to take the

complementary route to full-scale Reynolds number testing of

increasing tunnel size. The Propeller Research Tunnel (PRT), which

went into operation in July 1927, had an open jet test section 20 ft in

diameter and a stream velocity of 110 mph. The tunnel was used

mainly for tests on full-scale propellers, mounted in the fuselages of

real aircraft and driven by real engines. The propellers were full size,

running at their operational rotational speed and hence at virtually

full-scale Reynolds number. Many important advances came from

the ability given by this tunnel to test real hardware under realistic

aerodynamic conditions, including the development of the NACA

cowl for air-cooled engines, and led to NACA making the case for a

tunnel in which complete full-scale aircraft could be tested. Design

work on the Full-Scale Tunnel (FST) began in 1929 and the tunnel

began operations in spring 1931. It had an open jet test section of

30 ft60 ft (9.1 m18.3 m) and was driven by two 4000 hp (total

6 MW) motors, giving it a speed range of 25118 mph and a

maximum characteristic Reynolds number of 4.7 million. It played

a key role in US aircraft development in the 1930s and 1940s andremained in service for until 1995.

In 1939 another high Reynolds number tunnel came into opera-

tion at Langley, again with a drive of 8000 hp. This was the 19 ft

pressure tunnel, the first attempt anywhere to combine large scale

with high pressure. With a maximum pressure of 2.5 atm and a

maximum speed of 300 mph, its characteristic Reynolds number

was 11.9 million which enabled models of fighter and twin-engine

bomber aircraft to be tested at or near full-scale Reynolds number.

The advances in the USA were followed in Europe, both the UK

(Farnborough, 1934) and Germany (Braunschweig, 1940) building

8 m diameter tunnels in which, as in the NACA Propeller Research

Tunnel of 1927, full-scale propellers could be tested installed on an

aircraft. In France (Chalais-Meudon, 1934) a large tunnel with an

elliptical test section 16 m8 m was built. It was of the Eiffel type,with the open air rather than a hangar as the return circuit. Its

characteristic Reynolds number was 3.4 million. These and other,

smaller facilities played a part in enabling the respective national

industries to develop aircraft that would perform satisfactorily at

flight Reynolds numbers. All could be classed, however, as low-

speed tunnels, limited to testing aircraft at flight speeds at which

the flow around the aircraft could be treated as incompressible.

3.2. The significance of compressibility and the first high-speed

tunnels

As flight Mach numbers increase, the significance of compres-

sibility the local variation in air density caused by the passage of

the aircraft increases. Our insight into the behaviour of

Fig. 12. NACA Variable Density Tunnel (VDT) 1923.

Source:Ref. [13]p. 302.

7 We now know that free stream turbulence can also affect the development

of the turbulent boundary layer, as was recognised in specifying the turbulence

requirement for ETW (paper 4 in[35]).



9/50

compressible flows began in 1816 with Laplaces correction of

Newtons theory for the speed of sound which recognised the

importance of the ratio of the specific heats g. Later in the centuryEarnshaw and Riemann (in Gottingen) studied waves of finite

amplitude in a compressible fluid, Rankine and Hugoniot formu-

lated the equations for a plane shock wave, Ernst Mach photo-

graphed shock waves and de Laval patented the convergent-

divergent nozzle for generating supersonic flow.

In Gottingen in 1908 Theodor Meyer, a doctoral student ofPrandtl, submitted a thesis in which the key relationships for

supersonic flow were developed, including the formulation of the

expansion fan in supersonic flow around a sharp corner (the

PrandtlMeyer expansion) and the equations for an oblique shock

wave. In his lecture in 1920 on the Gottingen wind tunnel, Prandtl

referred to the propeller drive system under development on

which a propeller of 1 m diameter will probably be driven at a

speed as high as 5000 rpm for the purpose of studying the

influences of the compressibility of the air. At the time, however,

the typical maximum aircraft speed was around Mach 0.15 and

compressibility, whilst it might affect the flow around a propeller

blade, had no significant effect on the flow around an aircraft.

In the years following World War I, interest in the effect of

compressibility on flow around propellers spread. However, with

wind tunnel drive power increasing as the test section velocity to

the power three, the power required to drive a continuous flow

tunnel of any size up to the speeds of propeller tips was too great

to be contemplated. In the USA NACA, after funding various small-

scale aerofoil tests in high speed jets at other sites, began in 1927

to design its own, small, high-speed tunnel. This was an inter-

mittent tunnel with an 11 in diameter test section, supplied by

atmospheric air drawn through the test section by a downstream

ejector (the same principle also underlay one of the candidate

drive systems considered for ETW). The air to drive the ejector

was supplied from the pressure vessel that was the outer shell of

the VDT, its capacity being sufficient for a run of approximately

1 min. The Reynolds number on an aerofoil with a chord of 2 in

was approximately 0.8 million.

Results from the tunnel [17]revealed a sharp rise in aerofoildrag as Mach number increased towards unity. The success of the

11-in high speed tunnel prompted the design of a larger version,

the 24-in high speed tunnel, and inspired the UK to build a 12 in

diameter tunnel of the same design which also mimicked the

NACA tunnel by drawing the air for its ejector drive from the NPL

Compressed Air Tunnel at Teddington. Both tunnels came into

operation in 1934 and made valuable contributions to improving

propeller performance over the next decade. In Germany the first

significant high-speed tunnel, built in Gottingen in the late 1930s,

was also a short-duration intermittent tunnel supplied by atmo-

spheric air but with the refinement that the air was dried by being

drawn through a silica-gel filter. It was based on a concept first

set out by Prandtl in 1912 of drawing air through the tunnel into a

large vacuum vessel. It had an open-jet test section and wasequipped with interchangeable nozzles, 11 cm 11 cm for sub-

sonic testing (0.5oMo1.0), 11 cm13 cm for supersonic testing

(the tunnel had a range of Laval nozzles, 1.2oMo3.2). It was in

this tunnel in the autumn of 1939 that Ludwieg (who in 1955

invented the Ludwieg Tube drive system considered for ETW)

made the first measurements on a swept wing at high subsonic

and supersonic speeds. He thereby confirmed the validity of the

swept wing concept for supersonic aircraft that had been advanced

by Busemann of Gottingen at the Volta Conference in Rome in 1935.

The limitations of the intermittent high-speed tunnels at Langley,

small model size and limited testing time, led in 1933 to NACA

beginning the design of a large continuous running tunnel. This, the

Langley 8-ft high-speed tunnel, was completed in March 1936. It

was the first, and for 5 years the only, large high-speed wind tunnel

in the world. Built of reinforced concrete, driven by an 8000 hp

motor and with a maximum Mach number of 0.75 in a test section

8 ft in diameter, it had a maximum characteristic Reynolds number

of 3 million and played an important part in the development of US

combat aircraft in World War II. A particularly valuable contribution

was the solution of the problem of severe buffeting and loss of

control that affected the Lockheed P38 fighter in a steep dive, arising

from shock wave oscillation on the wing at high subsonic speeds.

The problem was cured by the development in the tunnel of a diveflap on the lower surface of the wing.

In the early 1940s other large high-speed tunnels came into

operation. New 16ft tunnels were built at NACA Langley and NACA

Ames, in the UK a 10 ft7 ft high-speed tunnel was built at

Farnborough and in Germany three high speed tunnels with test

section diameters in the range 2.7 to 3.0 m were built. By 1945 there

were also a number of medium-sized supersonic tunnels in

Germany, including a continuous running tunnel of 94 cm94cm

test section at Braunschweig which covered both the subsonic range

and supersonic Mach numbers between 1.1 and 1.8. The most

ambitious German project, launched in 1940, was the 8 m diameter

high-speed tunnel to be built in the Austrian Alps in the Otztal. This

was to be the largest high-speed tunnel in the world. It was not

completed when World War II ended and the components were

thereafter transferred to Modane, in the French Alps, where it became

the major facility at the newly created ONERA test centre. Its original

specification was for a maximum speed of 300 m/s and a character-

istic Reynolds number of 8.5 million. This called for a drive power of

76 MW, a very severe demand which was to be met by locating the

tunnel in a mountain valley and driving a pair of contra-rotating fans

by a pair of water turbines (Pelton wheels) supplied from a mountain

reservoir 530 m above. The rebuilding at the ONERA Modane-Avrieux

centre was completed and the tunnel went into service as S1MA in

1952. The drive was, as originally conceived, by Pelton wheels

supplied from a reservoir in the mountains above. This has been

perhaps the most extreme example of the location of a major tunnel

being determined by its power requirement.

4. The great period of wind-tunnel building 19451959

In the years immediately following the end of World War II

there was a surge forward in planning new wind tunnels, driven

partly by the realisation of the advances in both wind tunnel and

aircraft design made in Germany during the war. After the war

some of the German tunnels were dismantled and re-built in the

USA, France and Britain and many leading German aerodynami-

cists were recruited into the government laboratories, often to

continue research in the fields in which they had been working

previously. Dietrich Kuchemann and several of his colleagues

came to RAE Farnborough at that time.

Even before the end of the war, German and British jetpropelled aircraft were in operational service and the potential

for future supersonic aircraft was evident. In 1945 NACA, which

already had pilot supersonic tunnels at Langley and Ames, set in

hand the design of three large, continuous operation supersonic

wind tunnels, a 4 ft4 ft tunnel at Langley, a 6 ft 6 ft tunnel at

Ames and, at the Lewis Flight Propulsion Laboratory at Cleveland,

Ohio, a tunnel for jet engine testing with a test section measuring

8 ft 6 ft. These all had substantial power requirements, the

Langley tunnel being limited to a maximum operating pressure

of 0.25 atm because there was only 6000 hp available to drive it.

In contrast, the propulsion tunnel at Lewis had a drive power of

87,000 hp. In the UK, also before the end of the war, it had been

decided to build a major government wind-tunnel and flight-test

centre. This was to include a number of high-speed wind tunnels



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and the availability of substantial electricity supplies was an

important factor in the decision to locate the centre at Bedford.

The weakness in these plans was that in 1945 there was no

credible way of testing an aircraft model of realistic size at Mach

numbers greater than 0.9 in the wind tunnels of the time, because

of the flow choking in the vicinity of the model. The problem of

wall interference, the influence of the wind tunnel wall on the

flow field around the model, had been known for many years and

theoretical treatments of the boundary effects for both closed andopen jet test sections, and the corrections to apply to test results

for these effects, had been developed for low-speed flow. In the

1940s there was work on wall corrections for high-speed subsonic

flow but the methods were not applicable at Mach numbers close

to unity. However, it was well known that the interference effects

from solid walls and open jet boundaries were of opposite sign.

Starting from this, Wright at NACA Langley [18] developed a

theoretical model of a tunnel with longitudinal slots in which the

opposing interference effects of the solid and open sections of the

wall cancelled each other to produce, ideally, an interference-free

flow.8 Wrights work in 1946 led to the construction of a pilot

tunnel with a 12 inch slotted test section. This was a success,

showing much less wall interference and enabling Mach number

to be increased progressively through Mach 1 to low supersonic

speeds simply by increasing fan speed.

The result was a decision by NACA to install slotted walls in

both the 8 and 16 ft high-speed tunnels at Langley. The 8 ft

tunnel, which in February 1945 had had its drive power increased

from 8000 to 16,000 hp to give it an empty tunnel Mach number

of 1.0, was the first to be modified. It became operational as a

transonic tunnel in early 1950 and was followed shortly by the

re-powered 16 ft tunnel. The 8 ft tunnel was the facility in which the

transonic drag rise problem of the first generation of supersonic

fighters was identified and, through Whitcombs development of the

Area Rule,9 was solved. The Convair F102 supersonic interceptor was

the first aircraft to encounter this problem. Although powered by the

worlds most powerful jet engine of the time, tests in the 8 ft tunnel

indicated, and flight tests on the prototype confirmed, that it could

not go supersonic in level flight. Area ruling, narrowing the fuselage inplaces and adding bulges ahead and aft of the waist, overcame the

problem and saved the project.

Up to this point the large high-speed wind tunnels had relied on

air exchange between the tunnel and the outside air to remove the

motor power that is put in through the fan. The original 8 ft tunnel,

with 8000 hp being put into the fan, required to exchange about 1% of

its airflow with the outside atmosphere to maintain tunnel tempera-

ture at an acceptable level. With its doubled power, needed to

overcome the increased losses caused by the slotted test section

and to drive the flow to higher Mach numbers, the required exchange

rate with the outside air doubled. In the summer, when the humidity

of the outside air at Langley was invariably high, humidity within the

tunnel was similarly high and the temperature drop in the

acceleration to high speeds caused dense fog in the test section,

water droplets interfering with instrumentation and a deterioration

in tunnel flow quality. Langley quickly put in hand the construction of

a new tunnel with slotted walls, the 8-ft transonic pressure tunnel,

which removed the fan power via a water-cooled heat exchanger and,

being a sealed tunnel, avoided the problems of moist air. It could

operate at up to 2 atm pressure and had a high contraction ratio plus

screens in the settling chamber to provide a high flow quality in the

test section. It went into operation in 1953 and was the first facility toincorporate the essential features of a modern transonic tunnel.

Following the early experiments with slotted walls at Langley

there was work at Ames and the Cornell Aeronautical Laboratory to

explore the alternative of porous and perforated walls (walls with a

mesh of small holes) as a possible means of reducing the strength of

the reflection at the wall of the bow shock wave from the model. This

was significant only at Mach numbers around 1.0, where the reflected

shock could strike the rear of the model, but for military aircraft the

near-sonic Mach number range was a critical one. It was found that

perforated walls did indeed cause less interference than slotted walls

in the low supersonic regime effectively eliminating the reflected

shock and it was decided therefore to convert the 16 ft diameter

high-speed tunnel at Ames, which had been operating since 1941,

into a 14 ft14 ft transonic tunnel with perforated walls. The

increased test section drag, caused by the displacement of air into

the plenum chamber by the model and its subsequent return to the

tunnel stream with much reduced total pressure, together with an

increase in maximum test section Mach number, required the drive

power to be quadrupled, from 27,000 to 110,000 hp (82 MW). The

Ames 14 ft transonic tunnel began operation in 1955.

During this hectic period of tunnel development in the USA there

was also intense activity in Europe to create a new generation of

wind tunnels. In 1952, France had transferred the components of the

German 8 m high-speed tunnel from the Otztal to Modane and had

completed the construction of the tunnel as S1MA, thereby inaugu-

rating the ONERA Modane-Avrieux centre. In the UK, work was in

progress on a new flight and wind tunnel test centre near Bedford, a

very ambitious government project known at the time as the

National Aeronautical Establishment (NAE). By 1952 the first windtunnel was already in operation there, a 3 ft 3 ft supersonic tunnel

driven by plant originally used in the 94 cm94 cm supersonic

tunnel at Braunschweig mentioned above.10 A new 8 ft8 ft sub-

sonic/supersonic tunnel was also under construction to add to the

capability provided by the 10 ft7 ft tunnel at RAE Farnborough,

which was the only large high-speed tunnel in the UK at

the time. Also in the UK, in January 1952, the Aircraft Research

Association (ARA) was founded by 14 aircraft and engine compa-

nies with the specific aim of building a large high speed wind

tunnel for industrial project development work. In the Nether-

lands, after a hiatus of 28 months caused by a funding crisis, work

began again on a high-speed tunnel for the government aero-

nautical laboratory, the then NLL, in Amsterdam. Because the

power requirements for this tunnel were high relative to thecapacity of the local electricity supply, the 20 MW required to

drive its electric motor was supplied from a battery of five oil-

fired steam turbine power plants acquired as war surplus from US

navy destroyers.

At the time of this activity, the work at NACA to develop a

transonic test section was classified and work on the high speed

tunnels in Europe was proceeding without the benefit of the

US test section design knowledge. There is an account [21] by

8 The idea of a test section with a combination of solid wall and free air

boundaries has been attributed[19]to Prandtl (Gottingen) and Glauert (Farnbor-

ough) in the 1920s and to work by Wieselsberger of Gottingen in 1942 and Ferri in

Italy. Even so, there seems no doubt that Wright developed his theoretical model

independently and was the first to put it to the test of experiment. To add a

footnote to a footnote, it is worth recording that Glauert, an Englishman who was

a predecessor of Kuchemann as Head of Aerodynamics Department at RAE, was

killed on Saturday 4 August 1934 in an accident on the edge of the Farnborough

airfield when army engineers were using explosive charges to remove a tree

stump. The first author recalls Kuchemann recounting how, when the news

reached Gottingen, Prandtl called the laboratory staff together and said Gentle-

men, Glauert has been killed; we will do no work today and sent them home.9 Although many German aerodynamicists had emigrated to the US at the end

of WWII, it appears that Whitcomb discovered the area rule for himself, unaware

that it had been discovered by Frenzl in Germany in 1943[20]and was covered by

Junkers patent 932410 of 21 March 1944.

10 From 1971 to 1973 this tunnel was part of the first authors responsibilities.

Each year, the annual inspection of the compressors brought from Germany after

the war led to long deliberations as to whether the fatigue cracks in the casings

were getting worse. Finally, because of the cracks, the tunnel was taken out of

service in 1983.



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von Karman, then Chairman of AGARD, of how he helped to

ensure that the NLL tunnel was finally designed as a transonic

tunnel, with a slotted test section inside a plenum chamber,

rather than as a high-speed tunnel with solid walls. By the time

the design of the ARA Transonic Wind Tunnel was finalised,

Europe had been given access to the US advances in transonic

test section design and it was decided to adopt the Ames model of

perforated test section walls for the ARA tunnel. By 1954 the

original UK plans for a separate National Aeronautical Establishmenthad been drastically scaled down and the test centre had been

absorbed within the Royal Aircraft Establishment. The NAE became

RAE Bedford. The test section of its supersonic 3 ft3 ft tunnel was

adapted to accommodate slotted walls and in 1956 the former

10 ft7 ft high-speed tunnel at Farnborough returned to service

with slotted walls installed to convert it into the RAE 8 ft6 ft

transonic tunnel. In 1956 the ARA Transonic Wind Tunnel (TWT) ran

for the first time. It had a 9 ft8 ft test section with a flexible

supersonic nozzle ahead of the perforated test section and was

powered by a combination of a 25,000 hp motor driving the main

fan and a separate 14,000 hp motor driving an auxiliary compressor.

The latter was used at transonic and supersonic speeds to draw the

test section boundary layer air out through the perforated walls,

thereby reducing pressure losses in the diffuser and enabling the two-

stage fan to drive the tunnel up to Mach 1.4.11 In 1957 ONERA

brought the S2MA transonic-supersonic tunnel into operation in

Modane. This had separate test sections for transonic and supersonic

testing, the transonic section measuring 1.75 m1.77 m, and, as

with S1MA, it was powered by Pelton wheels. Also in 1957, the

8 ft8 ft tunnel at RAE Bedford was commissioned. This had a drive

power of 60 MW and a flexible supersonic nozzle with solid walls

through the test section which gave it a Mach number range from

low subsonic up to 2.812 but excluding the strictly transonic range. In

1959 the high speed tunnel at NLL (now NLR) came into operation

with a slotted 2 m1.6 m test section and a 20,000 hp drive. This

was the last new facility of its kind; the suite of the main post-war

transonic tunnels in Europe was now complete.

While Europe was building its large transonic tunnels in the

1950s, the USA was doing likewise. In 1949 the US Congress passedthe Unitary Wind Tunnel Plan Act and the Air Engineering Develop-

ment Center Act. The Unitary Plan embraced NACA, the USAF,

industry and universities and an early draught envisaged 33 large

transonic, supersonic and hypersonic wind tunnels costing almost $1

billion. As with the original UK plan for the National Aeronautical

Establishment at Bedford, budget realities resulted in a final plan of

more modest scale. Even so, the USAF Air Engineering Development

Center, now the Arnold Engineering Development Center (AEDC),

established in 1951 at Tullahoma Tennessee in order to be close to

the abundant hydroelectric power available from the Tennessee

Valley Authority, was a massive undertaking. It included two

16 ft16 ft wind tunnels to cover the Mach number range from

0.2 to 4.74. The transonic tunnel, with perforated walls, first ran in

1956 and since then has played a key role in all US military aircraftdevelopment. The other important transonic facility created under

this legislation was the NACA Ames transonic tunnel, with a slotted

test section 11 ft11 ft. This was part of the Ames Unitary Plan Wind

Tunnel Complex, which comprised the transonic tunnel and two

smaller supersonic tunnels, all linked to a set of motors and

compressors with an installed power of 180,000 hp. The Ames 11 ft

tunnel went into service in 1957. Other large high-speed tunnels that

were converted to transonic tunnels in the 1950s included the Boeing

8 ft12 ft tunnel in Seattle, the Cornell Aeronautical Laboratory (now

Calspan) 12 ft high speed tunnel at Buffalo, converted to a test section

8 ft8 ft in 1956, and the high speed tunnel at the Naval Surface

Warfare Center in Carderoc, Maryland, converted to a transonic test

section 10 ft7 ft in 1958.13

In addition to the large, fan-driven wind tunnels in Europe and

the USA there were also many smaller fan-driven or blow-downtransonic and supersonic tunnels built during that period, pri-

marily in industry but some also at universities. In his book on

transonic wind tunnels [22], published by AGARD in 1961,

Gothert lists 19 transonic tunnels in Europe and 30 in the USA.

5. Emergence of the need for higher Reynolds number

19591968

By 1960 the NATO nations had at their disposal an impressive

array of transonic and supersonic wind tunnels suitable for aircraft

development testing (there had also been a large programme of

wind tunnel building at TsAGI in Zhukovsky, near Moscow, but little

was known of that in the West at the time). Overall, the largest

tunnels in the USA had higher maximum Reynolds numbers thantheir European counterparts but the difference was not great. The

investment in these facilities had been substantial and there was a

feeling that, for transonic and supersonic testing, the job had been

done. On both sides of the Atlantic, the NATO nations now had the

wind tunnels they needed.

Gotherts book on transonic wind tunnels [22] sets out in

detail the level of understanding that had been reached in 15

years of intensive post-war development. In 1962 the Interna-

tional Union of Theoretical and Applied Mechanics (IUTAM) held a

Symposium Transsonicum in Aachen. In looking back on this in

1969, Kuchemann saw it as a meeting held at a time when many

of the researchers in transonic aerodynamics had already moved

to other fields, mainly space research. They had come back to the

meeting to present and sum up results which, in many cases, theyhad obtained long before. Since 1962, work in transonic aero-

dynamics had continued only at a relatively low level, the main

research being carried out by a few workers.

The priority given in the 1940s and the 1950s to transonic

aerodynamic research, and to the development of transonic wind

tunnels, had arisen mainly from the quest to develop supersonic

fighter aircraft. The transonic region had been important primarily as

one which the aircraft had to traverse controllably; once aerodynamic

knowledge and engine thrust had advanced to the point where that

hurdle could be cleared comfortably, which by 1960 they had, the

interest in transonics fell away. This was, however, only a temporary

fall in interest. In 1958 jet travel across the Atlantic began, the Comet

4 in September and the Boeing 707 in October. Both aircraft were

based on the late 1940s to the early 1950s aerodynamics but theirintroduction was followed by a rapid growth in air travel in the 1960s

and a demand from the airlines for larger and more efficient aircraft.

Also, in the spring of 1960, the US Air Force released Specific

Operational Requirement 182for a long-range freight aircraft, to which

Lockheed responded successfully with a large, turbofan-powered

swept-winged design, the Lockheed Model 300, subsequently desig-

nated C-141.

11 Many transonic tunnels use auxiliary suction to supplement the main fan

drive. Gothert[22]discusses the minimisation of total drive power by optimising

the balance between fan and auxiliary suction power.12 In 1970 the tunnel compressor was modified, reducing the top Mach

number to 2.5 in order to increase Reynolds number at high subsonic speeds to

approximately 8 million. The tunnel was taken out of service in 2002 and has since

been dismantled.

13 The structure of this tunnel and its drive fans a pair of cast steel contra-

rotating fans 19ft in diameter came from a 3 m high speed tunnel at Ottobrun

near Munich, an ambitious project that had not begun operation when the war

ended; the Carderoc tunnel went out of service in 1990 when one of the fans

suffered a catastrophic fatigue failure. The first author visited Carderoc shortly

after the failure and witnessed the devastation caused, even though the fan was

contained within a concrete shell.



12/50

As Kuchemann [23]noted, by the time of the AGARD Specia-

lists Meeting in Paris in September 1968 on the subject of

Transonic Aerodynamics there had been a general revival of

interest in the subject, with the participants from Industry stating

that the importance of continued technical advances in this field

cannot be overemphasised. The interest now was in the next

generation of transport aircraft, both civil and military. These

were subsonic aircraft with moderately swept wings on the upper

surfaces of which, at the cruise condition, there was an embeddedregion of supersonic flow terminated by a shock wave. Three

months before the meeting the Lockheed C-5A Galaxy had begun

flight testing and 10 days after the meeting the first Boeing 747

was rolled out. This was almost 5 years after the first flight of the

Lockheed C-141 Starlifter on 17 December 1963, the 60th anni-

versary of the Wright Brothers first powered flight, and it was the

aerodynamics of the C-141 wing that raised the concerns that led

finally to ETW. There is an excellent account of the emergence of

ETW from this starting point in the book The European Transonic

Wind Tunnel ETW A European Resource for the World of

Aeronautics [24]by Jan van der Bliek, former Director of NLR and

the original member for the Netherlands on the ETW Steering

Committee. There is inevitably appreciable overlap between that

book and this paper and in some places we have unashamedly

borrowed van der Blieks words. In general, however, we discuss

the technical issues more fully and the policy and political issues

less fully than he does.

In his 1961 book [22], Gothert made no reference to Reynolds

number but by then the test centres and the industry had between

them developed their test methods and their procedures for extra-

polating from wind tunnel to flight. In 1958 Braslow and Knox [25]

had published a method for designing boundary-layer tripsnarrow

bands of distributed roughness to cause transition from a laminar to

a turbulent boundary layer. These bands were fixed to the wind

tunnel model, close to the leading edges of lifting surfaces, around

the aircraft nose, etc., so as to ensure a turbulent boundary layer

over effectively the entire surface of the model. This simulated flight,

to the extent that the full-scale aircraft would also have a turbulent

layer all over, and the correction to quantities such as drag to allowfor the difference between the Reynolds numbers in wind tunnel

and flight could be made relatively straightforwardly on the basis of

the then current understanding of the turbulent boundary layer.

Individual test centres had their own preferred method of tripping

the boundary layer and each company had its own methodology for

extrapolating from wind tunnel to flight. There was an acknowl-

edged, accepted level of uncertainty in this process but, overall,

testing in the major wind tunnels of the time was considered to be a

satisfactory basis for the aerodynamic design of a new aircraft.

In the 1960s advances in wing design changed the situation.

The problem became apparent at high subsonic Mach numbers

where there is a region of supersonic flow terminated by a shock

wave on the wing upper surface. As either Mach number or lift is

increased, a bubble of shock-induced boundary layer separationforms at the foot of the shock. With further increase in Mach

number or lift coefficient the shock strength and the extent of the

bubble increases until there is a sudden drop in trailing edge

pressure and loss of lift. For the aerofoil designs used in the early

1960s, the characteristic behaviour was for the bubble to grow

slowly with shock strength, with the chordwise position of the

shock remaining relatively unchanged until a critical condition

was reached in which the bubble expanded suddenly to cause the

drop in lift. This behaviour was not particularly scale sensitive, as

explained by Pearcey et al. [26], and observed differences between

wind tunnel and flight were not great.

Advances in wing design aimed at reducing wing weight led to

increases in wing thickness and aerodynamic loading. The newer

designs had a longer region of supersonic flow on the upper surface,

followed by a steeper pressure rise towards the trailing edge. It was

found that, at wind tunnel Reynolds numbers, the effect of increas-

ing Mach number or lift on the more recent aerofoil designs was to

generate both a bubble separation beneath the shock wave and a

separation at the trailing edge, the interaction between the two

determining the chordwise shock position. This feature resulted in

the shock position being sensitive to Reynolds number.

The first manifestation of this effect that was of practical impor-

tance was on the C-141. Designed on the basis of wind tunnel testing,its flight test results were sufficiently at variance with the wind tunnel

to put the viability of the project at risk for a while. Figs. 13 and 14,

from Loving [27], show wing pressure distributions on the wing upper

surface in wind tunnel and flight at Mach 0.75 and 0.85, respectively.

The wind tunnel tests followed the standard practice of the time, with

transition fixed near the leading edge. At a Mach number of 0.75,

where the flow of the upper surface was subcritical (i.e. Mo1.0

everywhere), the difference between wind tunnel and flight was

relatively small and consistent with previous experience. At a Mach

number of 0.85, however, the flow over the upper surface was

supercritical, reaching a peak Mach number of 1.32 in the flight case,

and the chordwise positions of the shock waves terminating the

supersonic region differed by about 20% of chord between tunnel and

flight. The result was a nose-down pitching moment in flight that was

appreciably higher than in the tunnel and, in consequence, the

tailplane downloads needed to trim the aircraft were appreciably

Fig. 13. Comparison between wing upper surface pressure distributions in wind

tunnel and flight results for C-141 aircraft at subcritical conditions [27].

Source:NASA TN D-3580,Fig. 1.

Fig. 14. Comparison between wing upper surface pressure distributions in wind

tunnel and flight results for C-141 aircraft at supercritical conditions [27].

Source:NASA TN D-3580,Fig. 2.



13/50

higher than had been calculated on the basis of the tunnel tests. In the

event, a complete re-stressing of the aircraft was needed before it

could be decided that the project would meet its design requirements.

Lovings investigations[27]of the tunnel-to-flight discrepancy on

the C-141 were carried out in the NASA Langley 8 ft pressure tunnel.

They revealed that the tunnel results were strongly dependent on the

chordwise location of the transition strip. Loving found (Fig. 15) that

as the trip was moved rearwards, the upper surface pressure

distribution increasingly approached that in flight. With the tip

removed completely, to allow natural transition, the position of the

terminal shock was essentially the same in tunnel and flight. Black-

well[28]followed Lovings work with experiments in the Langley 8 ft

pressure tunnel on a large two-dimensional aerofoil that could be

tested at chord Reynolds numbers typical of tunnel tests on a

complete three-dimensional model (3.0 m

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